JPH0220500A - Transition control system for space-machine attitude controller - Google Patents

Abstract

PURPOSE: To minimize the transition time by determining a horizontal angular momentum from roll angle, yaw angle and the respective rates at an interval of 1/2 of the nutation period of a spacecraft, and further calculating a firing time making the momentum zero. CONSTITUTION: An error computer 40 calculates the horizontal angular momentum vector consisting of a roll component and a yaw component of a spacecraft by the roll angle signal from an earth sensor 36 and the yaw signal, a yaw rate signal, and a roll rate signal by a rate assembly 34 at an interval of substantially one-half nutation period of a spacecraft dynamics 32. This double component vector signal is passed through noise filter means 52, 54 to calculate the firing time length making the horizontal angular momentum zero with prescribed constants 60, 62, and roll and yaw thrusters 70, 72 are fired through a timer 68 to perform a movement for reducing the nutation amplitude. Thus, the nutation angle and transition time can be minimized.

Description

DETAILED DESCRIPTION OF THE INVENTION [Field of Industrial Application] The present invention relates to an attitude control system for a three-axis stabilized spacecraft. More particularly, the present invention relates to controlling the transition from thruster control to momentum wheel control in a bias momentum three-axis stabilized stationary spacecraft.

[Prior Art] One of the tasks in a spacecraft with a biased momentum attitude control system is to perform high-authority operations at relatively high angular velocities, such as those used for station-keeping maneuvers. ) control mode that relies on a biased momentum attitude control system, e.g. using a momentum wheel.
authority) control of transition to control mode. The transition from station-keeping mode to on-orbit mode is due to the high angular velocity before the transition, which is
(nutation) and must be performed carefully to avoid saturation of attitude control and possible loss of attitude control resulting in low authority on-orbit control mode.

The desired control accuracy for a three-axis stabilized spacecraft is between 0.01 and 0.05 degrees, not counting misalignment and pointing errors due to structural or thermal distortion. However, the typical uncorrected attitude error associated with station-keeping maneuvers is on the order of 25 times larger, ie, a nutrition circle with a radius of 0.25 degrees to 1.25 degrees. In order to achieve a stable on-orbit mode, the nutation amplitude must be rapidly reduced, preferably to zero, before control is transferred to the momentum wheel-based on-orbit controller. .

Problems to be Solved by the Invention Most of the New Taylor 3 correction systems are oven loop control systems that are based on calculations rather than measurements of critical parameters. The following patents relating to nutrition control systems were revealed in a search of the U.S. Patent Office public records.

U.S. Pat. No. 3,866,025 discloses a spacecraft attitude control system that provides thruster firing to align the total angular momentum vector with the desired orbit normal vector while simultaneously adjusting the orbit. The invention uses commonly available sensors and thruster systems. However, this control system attempts to perform trajectory adjustment maneuvers in a manner that attempts to minimize the accumulation of nutrients during the maneuver, and does not provide a mechanism for correction following the maneuver. The angular velocity is calculated rather than directly sensed. Moreover, this patent does not mention the specific transition attitude control technique of the present invention.

Other patents identified that contain additional information on common issues such as nutrition attenuation, correction in spacecraft systems, etc. are:

Mata-ku Tateoka 1 U 3, 643.897 Ji 3 Inson, Junior 3,
937.423 Johansen 3.944472
Belaka 3! 184,071 Fleming 3.997.
137 Phillips 4.023.752
Bistiner et al. 4.174.819 Bliss, Dahl et al. 4.370,716 Amu 4
．． 386,750 Hofmann 4.521,855 Lehner et al. U.S. Pat. No. 3,643,897 relates to a spin-stabilized spacecraft, not a three-axis stabilized spacecraft as the present invention does. Although this patent attempts to hold the nutation angle to a limited value, there is no suggestion as to how to reduce the nutation to zero.

No. 3,937,423 discloses an attitude control system for a three-axis stabilized spacecraft. However, this patent does not disclose any mechanism for sensing yaw angle, yaw angular velocity, and roll angular velocity.

No. 3,944,172 relates to a spacecraft without internal momentum wheel control and is therefore unaffected by nutrition.

No. 3,984,071 relates to damping a nutrition circle by firing two thrusters based on sensing only the spacecraft roll angle. There is no means to sense roll, yaw, and yaw angular velocity. As a result, stability cannot be guaranteed.

U.S. Pat. No. 3.997.137, 4.023.
No. 752 and No. 4.370.716 relate to spin-stabilized spacecraft, not three-axis stabilized spacecraft, and are therefore not relevant to the present invention.

Although the aforementioned U.S. Pat. No. 4,174,819 relates to attitude control of a three-axis stabilized spacecraft, it does not address the problem of nutrition control and the problem of transition from thruster control mode to wheel control mode. Also, there are no sensors for yaw angle or yaw 1-speed.

[Means for Solving the Problems] According to the present invention, in a three-axis stabilized spacecraft that uses a momentum wheel fixed to the internal spacecraft body as an attitude stabilization device, the nutrition angle and transition time are minimized. A method is provided for controlling the transition from thruster control to momentum wheel control. This method uses direct, substantially real-time measurements of momentum components and angular velocity to generate paired thrust impulses that are synchronized with the spacecraft's nutrition period. In the ideal case with no measurement errors or thruster pulse errors, the nutation is zeroed out at 1/2 of the nutation period. However, since errors always exist, the controller uses a deadbeat principle in a closed-loop real-time feedback system. In a preferred embodiment, the impulsive thruster firing commands in roll and yaw are simultaneous with a time interval of I/2 nietation period, and the spacecraft lateral momentum is equal to or less than three consecutive pairs of thruster pulses. This has the effect that the ignition is reduced to an acceptably low level.

The present invention automatically reduces the angular momentum for a desired on-orbit condition to a level that can be handled by a low power mentum wheel control device with minimal transient overshading.

The present invention will be better understood by the following detailed description taken in conjunction with the accompanying drawings. The description herein is based on a symmetric spacecraft, ie, a spacecraft with equal yaw and roll inertia, for ease of explanation. However, the principles of the present invention can be applied to any spacecraft.

EXAMPLE FIG. 1 is a vector diagram 10 of a spacecraft illustrating a nutrition circle, typically an ellipse, about a set of lateral (roll/yaw) momentum axes. For simplicity, the body and structures of the spacecraft itself are not shown. One of the objectives of the invention is to minimize the nutrition of the spacecraft, ideally reducing the nutrition to zero. It is to let FIG. 1 shows the effect of suitable angular impulse correction according to the invention.

(Some basic principles will be helpful in understanding the invention.

First of all, the change in spacecraft momentum, vector ΔH, is equal to the product of the thruster torque, vector r, and the ignition time t. The vector ΔH can be decomposed into a roll component on the X axis and a yaw component on the Z axis, and as a result, the vector equation ΔH=TΔt can be easily expanded as follows.

8H* = T Δt ΔHv ”” T vΔt ΔI (Eng = T, Δt where Δt is the ignition time. Since the pitch ignition is not included, the y component is O, and therefore the pitch is Zero ignition time, which is negligible for the purpose, is 40 m5 to 50
On the other hand, the diatation period 17λ, that is, the time for the spacecraft body to completely nutate the nutation circle, is about 150 to 300 seconds or more. Long nutation periods allow approximations related to impulsive ignition. The nutrition period is a function of the inertia Ise of the spacecraft and the wheel motion @H of the spacecraft somentum wheel. Ignition of the thruster changes the center of the two-knee joint and roll angle φ
is proportional to the yaw momentum H3, i.e. φ. =-ΔH
z/Hwx, and the yaw angle ψ is the roll momentum H,I
Proportional to, Δψ0=ΔH, /H.

Pitch control is independent of roll and yaw, so only roll and yaw need be considered.

The vector diagram 10 showing the roll and yaw angle axes represents the plane of the nutrition circle.

In particular, one of the objects of the invention is to reduce the nutation and the corresponding nutation circle 12 with its center 14 at C0 to O. According to the invention, this measures four independent variables: roll angle, roll rate, yaw angle and yaw rate, and then uses these measured quantities to generate thruster & impulse thrust; It is necessary to move the center 14 of the nutation by some momentum vector (Δh+=Δhx+Δh□) to a new center 16 of C1 closer to the origin, and then repeatedly move the center of the nutation, ideally to the origin. do. Actually the center of nutrition of nutrition circle 19 is 16°
exists, and in response to the actual change in the momentum vector Δh2, a third impulse Δhs or a subsequent (impulse According to the invention, these momentum changes are achieved by allowing simultaneous impulsive firing of the spacecraft's roll and yaw thrusters. , i.e.,
1=0 and t=π/λ. Here, 1/λ is the nutrition period. The nutrition period in a typical spacecraft is about 300 seconds. Ideally,
That is, in the absence of error, the firing of the second thruster at t=:W:/λ occurs at the origin, so the center of nutrition is at the origin, and the It will be done. In reality, three ignitions will turn the nutrition on.
It was found necessary to reduce the

FIG. 2 shows the configuration of a control system 30 according to the present invention.
The spacecraft is metaphorically represented by spacecraft dynamics 32. Spacecraft dynamics 32 is generated using the spacecraft motion as an output parameter, and the direction and orientation of this spacecraft motion is determined by the digital inch grating rate assembly (DIRA) 34, which is a gyroscope onboard the spacecraft, and by the Earth. Measured by sensor 36.

The roll signal φ is sent from the earth sensor 36 to the built-in error calculator 4.
0 herol signal line 38. Also,
DIRA provides as outputs a yaw signal ψ on a yaw signal line 42, a yaw rate signal gold on a yaw rate signal line 44, and a roll rate signal ↓ on a roll rate signal line 46 to an internal error calculator 40.

As will be described below, a built-in error calculator 40 calculates a predetermined value for roll error em and a predetermined value for yaw error 62 from these first measured parameters and transmits these signals to signal lines 48 and 50 or to noise filter means 52 and 54 via equivalents of . The noise filter means 52 and 54 are essentially low pass filters whose function is to filter high frequency rate noise from the roll and yaw error signals due to measurement errors and structural flexure effects in the spacecraft. 1st
The output signal of the second noise filter 52 is the roll error signal e8 on the signal line 56, and the output signal of the second noise filter 54 is the roll error signal e! on the signal line 58. It is. The two error signals are then constant values for these purposes: the roll inertia component divided by the lateral roll torque component l:X, and the yaw inertia component 2 divided by the lateral yaw torque The weight is given by the value 62 divided by the component child. Output signals 64 and 66 are timer 68
The timer 68 uses these signals to control the impulse period for each roll thruster 70 and yaw thruster 72. These thrusters 70
and 72 are torque Tc, (roll torque) and Tcx (
yaw torque) to the spacecraft, and the spacecraft dynamics 32 are changed accordingly. Ignition of the thruster causes the center of the new chishiran to change.

According to the invention, the supplied torque is the firing of paired or deadbeat thrusters with a duration and vector determined by the measured roll, yaw, roll rate, and yaw rate, resulting in lateral momentum. towards 0, and the center of the new chishiran to yaw and roll O
decrease towards. The error signals for roll and yaw errors are calculated from the measured pan by the following kinematic equations:

ez=(Φ+1/2ED) ex=-(↓-1/2Eψ) These equations, when properly filtered and weighted by a constant, give ideally provides enough information for all of the error corrections needed to yield a three-dimensional circle that passes through the origin. Assuming that the firing of the first thruster is calculated as above, the roll and yaw angles approach O after 1/2 a day period (P, in Figure 1),
Firing of the second set of thrusters at this time, calculated from the sensor data according to the above formula, results in a very small new chisi run circle. This process can be carried out by 1 if necessary.
It can be repeated with one or more additional ignitions. In practice, firing an additional set of roll/yaw thrusters after an additional 1/2 rotation period has been found to be very beneficial.

Pitch is not a factor in the Nutasigin correction since pitch control is independent of roll and yaw; thus, by knowing roll, yaw, roll rate and yaw rate instantly, and knowing the spacecraft's moment of inertia, The momentum of the spacecraft can be calculated in two dimensions and the momentum can be corrected.

Although the present invention has been described with respect to particular embodiments, other embodiments will be apparent to those skilled in the art from the description herein. May be based on measurements;
Alternatively, it can be exactly approximated by a constant. Similarly, when calculating the lateral momentum component, the different roll and yaw moments of inertia, as well as the product of inertia, if meaningful, may be taken into account. A third variation can occur if the control thruster has significant loss coupling torque, since this adversely affects the calculation of the thruster firing time. What calculations do you need? Depending on the complexity, digital processors may be preferable to analog electronics (circuits). Accordingly, the invention is not intended to be limited except as indicated by the claims.

Figure 1 shows the deadbeat impulse ignition of a three-axis stabilized spacecraft in which a circle representing a circle or an ellipse is shown for simplicity. A vector diagram illustrating a reduction in lateral momentum by control, and FIG. 2 is a configuration diagram of an embodiment of a spacecraft momentum control system according to the present invention.

10: Vector diagram of spacecraft 12.19: New technology circle 30: System (wholesale system 32: Spacecraft dynamics 34: Digital inch grating rate assembly (D I RA) 36: Earth sensor 40: Error Calculator 52.54: Noise filter means 60.62: Constant 68: Timer 70: Roll thruster 72: Yaw thruster

Claims (6)

[Claims] (1) In a three-axis stabilized spacecraft that uses a momentum wheel fixed to the internal spacecraft body as an attitude stabilization means and is characterized by a nutation period, thruster control is performed to minimize the nutation angle and transition time. The method for controlling the transition from to momentum wheel control is as follows: a) roll angle of the spacecraft, yaw angle of the spacecraft,
obtaining values representing the spacecraft roll rate and the spacecraft yaw rate, respectively; b) obtaining the spacecraft moment of inertia, the spacecraft roll angle, the spacecraft yaw angle, the spacecraft roll rate, and the spacecraft yaw rate; c) calculating a lateral angular momentum vector of the spacecraft consisting of a roll component and a yaw component from the values respectively represented; c) calculating an ignition time length for each of the set of attitude control thrusters; d) igniting the set of attitude control thrusters in accordance with the calculated ignition duration; e) repeating steps a) to d) at least once at intervals of substantially 1/2 of the nutrition period of the spacecraft. The transition control method according to claim 1, further comprising the step of: (2) attenuating spurious signals substantially higher than the nutrition frequency of the spacecraft. 2. The transition control method according to claim 1, further comprising the step of: (3) calculating the nutrition period from the measured rotational speed of the momentum wheel and the moment of inertia of the wheel and the spacecraft. 4. The method of claim 2, wherein the thrusters are arranged at right angles and the firing step includes simultaneous firing to control roll and yaw. (5) In a three-axis stabilized spacecraft that uses a momentum wheel fixed to the internal spacecraft body as an attitude stabilization means and is characterized by a nutation cycle, thruster control is performed to minimize the nutation angle and transition time. an apparatus for controlling the transition from to momentum wheel control, coupled to the sensor means for obtaining values representative of a spacecraft roll angle, a spacecraft yaw angle, a spacecraft roll rate, and a spacecraft yaw rate, respectively; and control means for controlling the ignition of the roll and yaw attitude control thrusters of the spacecraft at intervals of substantially 1/2 of the rotation period, the control means controlling the moment of inertia of the spacecraft and the spacecraft. a device for calculating a lateral angular momentum vector of a spacecraft consisting of a roll component and a yaw component from the values representing the roll angle of the spacecraft, the yaw angle of the spacecraft, the roll rate of the spacecraft, and the yaw rate of the spacecraft, respectively; an apparatus for calculating a firing time length for an attitude control thruster and producing a motion of the spacecraft that reduces the nutation amplitude to zero transverse angular momentum following thruster firing; and said roll according to said firing time length. and a device for igniting a yaw attitude control thruster; and a device for igniting a yaw attitude control thruster;
and a timing device for sequentially igniting the roll and yaw attitude control thrusters at intervals of 2. 6. Claim 5 further comprising filter means coupled between said sensor means and said control means, said filter means removing spurious signals substantially higher than the nutrition frequency of said spacecraft. The transition control device described.